The present invention relates to the use of nuclear fission to heat gases. It has applications in the field of deep space rocket propulsion, in particular.
If nuclear energy is presently competing with other methods on Earth, for deep space travel it has unique features which make it practically indispensable in order to realistically attain the long range goals of a manned exploration of Mars, of the outer planets' satellites and of the Asteroids.
Neutron induced fission is the preferred nuclear reaction for practical energy production, because of (1) its remarkable energetic yield (≈200 MeV) and (2) its ability of sustaining the reaction with the secondary neutrons.
The present state of the art of nuclear space propulsion, is represented for instance by the NERVA design (see “Nuclear thermal rockets: next step to space”, Aerospace America, June 1989, pages 16-29; or R. W. Bussard, et al. “Nuclear Rocket Propulsion”, McGraw-Hill, New York, 1958). In the NERVA design, the energy produced by the fission reaction is recovered in the form of high temperature heat from fuel rods. To ensure the heat flow from the fuel to the propellant gas, the temperature of the propellant gas must be somewhat less than the operating temperature of the fuel, in turn limited by the mechanical strength and the stability of the fuel rods at high temperatures. Furthermore, in order to ensure a good heat transfer to the propellant, the pressure of the high temperature gas must be very high, of the order of 150 bars.
The exhaust velocity vexh of a rocket engine is limited by the enthalpy and the final temperature of the energy-producing reaction, and it is proportional to the so-called specific impulse. The specific impulse is defined as Isp=vexh/g, where g=9.81 m.s−2 is the gravitational constant. It represents the duration over which a given mass of propellant can exercise a thrust force equal to its weight. Chemical rocket engines using liquid hydrogen and oxygen typically operate at exhaust temperatures of 3600° K under stoichiometric conditions, with an effective molecular weight of about A=11, which limits the specific impulse to about 450 s. NERVA type engines perform somewhat better than chemical engines and offer a higher specific impulse of about 950 s. The advantage, however, is mainly due to reductions in the effective molecular weight (the specific impulse is proportional to 1/√{square root over (A)}) due to the use of pure hydrogen gas (A=2 vs. A=11) rather than an increase in the exhaust temperature.
In fact, NERVA rockets are expected to operate at lower temperatures than chemical rockets (3,000° K) due, as already pointed out, to material limitations of the reactor core. The vast amount of energy potentially available through the fission process remains largely untapped due to (1) size constraints associated with the minimum critical mass required to sustain the fission chain reaction and (2) the difficulty of extracting heat at sufficiently high temperatures from the reactor. Notwithstanding, the NERVA engine is often cited as being so far the only realistic engine candidate so far for a manned trip to Mars.
The NERVA engine is basically a naked fast reactor, which represents a serious drawback of nuclear engines for space propulsion. Let us consider for instance the set of three NERVA engines as described in a recent NASA Report on a Mars mission (“Human Exploration of Mars: The Reference Mission of the NASA Mars Exploration Study Team”, (including Addendum V3.0, June 1998), NASA SP 6107, 1997). The installed power is near 1 GWatt and about 3.2×1019 fast neutrons/s will be expelled from the engines. The standard allowed dose of ≦10 n/cm2/s is attained only at an unshielded distance of 5,000 km.
In addition, the neutron leakage will also hamper the simultaneous operation of several NERVA engines nearby, as in the above-mentioned report. Indeed, a reactor—even if switched off by control bars—is still a sub-critical, multiplying device and it will produce power if irradiated by neutrons from the nearby engines. For instance, if a mere 1% of the neutrons from one engine hit the neighbour unit scrammed at k=0.99, the latter will swing into full power. If it is already on, the additional neutron contribution will be sufficient to bring it to prompt criticality. The coupled control system for mutually interfering reactors is, in our view, a true nightmare and unrealistic in a manned space mission.
For any engine to be used for interplanetary travel, the residual neutron flux outside the engine should be sufficiently low as to permit the operation of the engine not too far from the Space Station (ISS) which is considered as the main “docking point” for the interplanetary journey. In addition, the dose given to the crew should also be small compared to the inevitable dose from the cosmic ray background, which amounts to about 40 rad/y.
The potential features of several nuclear devices for a Space Propulsion engine beyond the potentialities of NERVA have been illustrated by several papers (T. Kammash, ed., “Fusion Energy in Space Propulsion”, AIAA Progress in Astron. And Aeron., Vol. 167, AIAA, N.Y., 1995; or N. R. Schulze, “The NASA-LEWIS Program on Fusion Energy for space Power and Propulsion”, Fusion Technology, 19-1, pages 11-28, 1991). They are mostly based on Fusion rather than Fission, primarily because this process permits the use of charged reaction products directly to heat up the exhausted gas to high temperatures in the form of a plasma.
Both inertial and magnetically confined Fusion have been extensively explored. The choice of Fusion as a reference has been driven by the obvious argument that the ionising reaction products, which are used to heat-up the propellant, are much easier to extract from a Magnetically Confined (MC) or an Inertial Fusion (IF) device.
However, huge fundamental and technological problems have so far hampered the realisation of a practical Fusion, energy producing device on Earth, and even more so in space. Fusion machines, especially MC, are necessarily very large devices, of very complex technology and hardly adapted to the conditions of a long interplanetary journey.
Another concept of nuclear propulsion, based on fission, is the so-called plasma core propulsion. We mention the coaxial flow system and the nuclear light bulb engines (see R. Ragsdale, et al., <<Gas Core Rocket Reactors—a New Look>>, NASA TM X-67823, 1971 ; and J. D. CLEMENT et al., <<Gas Core Reactor Technology>>, Reactor Technol. 13-3, 1970). In these devices the fissionable material (enriched 235U) is allowed to heat-up to plasma temperatures, up to 50,000° K, and its radiation is used to heat up the hydrogen gas. This is not a trivial task, since hydrogen and most of the other light gases are optically transparent at temperatures less than about 15,000° K, except to their own radiation (lines).
Typically, a coaxial flow plasma reactor for space propulsion is expected to operate at 6,000 MWatt power, producing Isp=4,000 s. The cavity diameter is about 4 m, the pressure ranges from 400 to 600 bars and the total weight is of the order of 500 tons. The critical mass is between 40 to 80 kg of 235U. It is not clear how such a mass could be brought at start-up from solid to the plasma state.
The nuclear light bulb concept, unlike the coaxial flow system provides for full containment of the fuel within a transparent, internally cooled wall configuration, thereby circumventing the problem of fuel mixing with the propellant with a consequent loss in the exhaust. The fissioning plasma is kept away from the transparent walls by a tangentially injected swirl flow of buffer gas, which is re-circulated, with the Uranium losses recovered and re-circulated in the plasma. Otherwise, the principle of operation is the same as the coaxial flow plasma reactor. Typical data for the nuclear light bulb engine are: power 4,600 MWatt; Isp=1,870 s; weight 35 tons, edge of the fuel temperature 5,000° K and pressure 500 bars.
These concepts have been investigated in detail, though no test has been made. It is expected to be a difficult technology, the main concern being the control of the criticality of the Uranium plasma. Indeed a change in the multiplication coefficient of ≦0.7%—if not compensated promptly by the control bars—would lead to a prompt criticality accident. For a thick fissile material and neglecting the effects of the reflector, the critical mass is proportional to the inverse of the squared density. In addition, cross sections and hence the critical mass are functions of the temperature. Furthermore, it is not clear how an effective and safe control system can be realised in view of the rapid motion of the inner core (fissionable plasma and surrounding gas) and the possible emergence of hydro-dynamical instabilities.
An object of the present invention is to propose an alternative way of heating gases by means of nuclear fission reactions, which is suitable for space propulsion applications.